Hybrid turbine airfoil components containing a ceramic material, in which detailed features of the components are formed of materials other than ceramic materials. The components include a first component formed of a ceramic-based material and a second component formed of a metallic material. The first component comprises an airfoil portion and a nub, and the second component is separately formed and attached to the first component by casting the metallic material around the nub of the first component. The second component includes a platform portion between the airfoil portion and the nub of the first component and a dovetail portion on the nub of the first component. Each of the platform and dovetail portions has at least one off-axis geometric feature that results in the second component having a more complex geometry than the first component.
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1. A turbine airfoil component of a gas turbine engine, the airfoil component being a turbine blade comprising:
a first subcomponent comprising an airfoil portion and a nub, the nub comprising a base that has outer surfaces, the outer surfaces of the base defining a cross-section of the base that is wider than a region of the nub adjacent the airfoil portion, the first subcomponent being formed of a ceramic-based material and the airfoil portion and the nub being a unitary piece of the ceramic-based material; and
a second subcomponent comprising a casting formed and attached to the first subcomponent by casting a molten metallic material in-situ around the nub of the first subcomponent, the second subcomponent comprising a platform portion between the airfoil portion and the nub of the first subcomponent and a dovetail portion on the nub of the first subcomponent and adapted for anchoring the airfoil component to a turbine disk, each of the platform portion and the dovetail portion having at least one off-axis geometric feature that results in the second subcomponent having a more complex geometry than the first subcomponent, the dovetail portion of the second subcomponent having:
a shank casing that completely encases the nub so that the region of the nub adjacent the airfoil portion is covered by a nonconformal portion of the shank casing and so that the base of the nub is entirely covered by a conformal portion of the shank casing having an exterior surface that conforms to the outer surfaces of the base, the off-axis geometric feature of the dovetail portion comprising a first set of oppositely-disposed tangs that are exclusively defined by the nonconformal portion of the shank casing and protrude from the region of the nub adjacent the airfoil portion, the off-axis geometric feature of the dovetail portion comprising a second set of oppositely-disposed tangs that are exclusively defined by the base of the nub and by the conformal portion of the shank casing thereon, the conformal portion of the shank casing defining a pressure face on each of the second set of oppositely-disposed tangs.
12. A turbine airfoil component of a gas turbine engine, the airfoil component being a blade comprising:
a first subcomponent comprising an airfoil portion and a nub that is located on the shank portion and has a base having outer surfaces, the outer surfaces of the base defining a cross-section that is wider than an immediately adjacent region of the nub adjacent the airfoil portion, the first subcomponent being a unitary piece of a ceramic-based material; and
a second subcomponent comprising a casting formed and attached to the first subcomponent by casting a molten metallic material in-situ around the base of the nub of the first subcomponent such the nub is entirely encased within the metallic material that forms the second subcomponent and the second subcomponent is retained on the first subcomponent by a compression fit on the wider cross-section of the nub, the second subcomponent comprising a platform portion between the airfoil portion and the nub of the first subcomponent and a dovetail portion on the nub of the first subcomponent and adapted for anchoring the airfoil component to a turbine disk, each of the platform portion and the dovetail portion having at least one off-axis geometric feature that results in the second subcomponent having a more complex geometry than the first subcomponent, the dovetail portion of the second subcomponent having:
a shank casing that completely encases the nub so that the region of the nub adjacent the airfoil portion is covered by a nonconformal portion of the shank casing and so that the base of the nub is entirely covered by a conformal portion of the shank casing having an exterior surface that conforms to the outer surfaces of the base, the off-axis geometric feature of the dovetail portion comprising a first set of oppositely-disposed tangs that are exclusively defined by the nonconformal portion of the shank casing and protrude from the region of the nub adjacent the airfoil portion, the off-axis geometric feature of the dovetail portion comprising a second set of oppositely-disposed tangs that are exclusively defined by the base of the nub and by the conformal portion of the shank casing thereon, the conformal portion of the shank casing defining a pressure face on each of the second set of oppositely-disposed tangs.
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The present invention generally relates to ceramic-based articles and processes for their production. More particularly, this invention is directed to ceramic-based articles produced to include metallic regions that define detailed features, for example, dovetails, shanks, platform features and tip shrouds of gas turbine airfoil components.
Higher operating temperatures for gas turbines are continuously sought in order to increase their efficiency. Though significant advances in high temperature capabilities have been achieved through formulation of iron, nickel and cobalt-base superalloys, alternative materials have been investigated. Ceramic materials are a notable example because their high temperature capabilities can significantly reduce cooling air requirements. As used herein, ceramic-based materials encompass homogeneous ceramic materials as well as ceramic matrix composite (CMC) materials. CMC materials generally comprise a ceramic fiber reinforcement material embedded in a ceramic matrix material. The reinforcement material may be discontinuous short fibers dispersed in the matrix material or continuous fibers or fiber bundles oriented within the matrix material. The reinforcement material serves as the load-bearing constituent of the CMC in the event of a matrix crack. In turn, the ceramic matrix protects the reinforcement material, maintains the orientation of its fibers, and serves to dissipate loads to the reinforcement material. Silicon-based composites, such as silicon carbide (SiC) as the matrix and/or reinforcement material, are of particular interest to high-temperature applications, for example, high-temperature components of gas turbines including aircraft gas turbine engines and land-based gas turbine engines used in the power-generating industry.
Continuous fiber reinforced ceramic composites (CFCC) are a type of CMC that offers light weight, high strength, and high stiffness for a variety of high temperature load-bearing applications, including shrouds, combustor liners, vanes (nozzles), blades (buckets), and other high-temperature components of gas turbines. A notable example of a CFCC has been developed by the General Electric Company under the name HiPerComp®, and contains continuous silicon carbide fibers in a matrix of silicon carbide and elemental silicon or a silicon alloy. SiC fibers have also been used as a reinforcement material for a variety of other ceramic matrix materials, including titanium carbide (TiC), silicon nitride (Si3N4), and alumina (Al2O3).
Examples of CMC materials and particularly SiC/Si—SiC (fiber/matrix) CFCC materials and processes are disclosed in U.S. Pat. Nos. 5,015,540, 5,330,854, 5,336,350, 5,628,938, 6,024,898, 6,258,737, 6,403,158, and 6,503,441, and U.S. Patent Application Publication No. 2004/0067316. One such process is known as “prepreg” melt-infiltration (MI), which in general terms entails the fabrication of CMCs using multiple prepreg layers, each in the form of a tape-like structure comprising the desired reinforcement material and a precursor of the CMC matrix material, as well as one or more binders and typically carbon or a carbon source. The prepreg must undergo processing (including firing) to convert the precursor to the desired ceramic. Prepregs for CFCC materials frequently comprise a two-dimensional fiber array comprising a single layer of unidirectionally-aligned tows impregnated with a matrix precursor to create a generally two-dimensional laminate.
For purposes of discussion, a low pressure turbine (LPT) blade 10 of a gas turbine engine is represented in
Because they are directly subjected to hot combustion gases during operation of the engine, the airfoil 18, platform 16 and tip shroud have very demanding material requirements. The platform 16 and blade tip shroud (if present) are further critical regions of a turbine blade in that they create the inner and outer flowpath surfaces for the hot gas path within the turbine section. In addition, the platform 16 creates a seal to prevent mixing of the hot combustion gases with lower temperature gases to which the shank 20, its dovetail 12 and the turbine disk are exposed, and the blade tip shroud is subjected to creep due to high strain loads and wear interactions between its seal tooth (if present) and the shroud surrounding the blade tips. The dovetail 12 is also a critical region in that it is subjected to wear and high loads resulting from its engagement with a dovetail slot and the high centrifugal loading generated by the blade 10.
Current state-of-the-art approaches for fabricating ceramic-based turbine blades have involved integrating the platform 16, dovetail 12, airfoil 18 and tip shroud (if present) as one piece during the manufacturing process, much like conventional investment casting techniques currently used to make metallic blades. However, the platform 16, dovetail 12, tangs 14 and tip shroud represent detailed geometric features of the blade 10 that pose substantial challenges to designing, manufacturing and integrating CMC components into an affordable, producible design for turbine applications. For example, the process of integrating a platform 16 and tip shroud with the airfoil 18 using CMC materials creates complexities in the design and manufacturing process, and can result in a process that can be too expensive to be economically practical. Furthermore, the platform 16, dovetail 12 and its tangs 14 have interface/support functions that can require structural interface capabilities that can be difficult to achieve with CMC materials. In addition, the low strain-to-failure capabilities of typical CMC materials and the possibility of undesirable wear interactions between tip shroud seal teeth and conventional shrouding materials pose additional challenges to implementing CMC materials in shrouded blade designs.
The present invention provides hybrid turbine airfoil components containing a ceramic material, in which detailed features of the components are formed of materials other than ceramic materials, yet result in a robust mechanical attachment of the ceramic and non-ceramic portions of the components.
According to a first aspect of the invention, a turbine airfoil component includes a first subcomponent formed of a ceramic-based material and a second subcomponent formed of a metallic material. The first subcomponent comprises an airfoil portion and a nub, and the second subcomponent is separately formed and attached to the first subcomponent by casting the metallic material around the nub of the first subcomponent. The second subcomponent includes a platform portion between the airfoil portion and the nub of the first subcomponent and a dovetail portion on the nub of the first subcomponent. Each of the platform and dovetail portions has at least one off-axis geometric feature that results in the second subcomponent having a more complex geometry than the first subcomponent.
In view of the above, it can be seen that a technical effect of this invention is the ability to produce certain portions of a turbine airfoil component from a ceramic-based material, while producing other portions of the component having intricate geometric details from materials that do not require the temperature capability of ceramic-based materials. The invention is particularly beneficial for applications in which the intricate geometric details formed of the non-ceramic material are interface/supporting features that require structural interface capabilities, and as a result of being fabricated from a non-ceramic material are not nearly as labor intensive or require the level of skilled labor that would be required if the entire component were fabricated from a ceramic-based material.
Other aspects and advantages of this invention will be better appreciated from the following detailed description.
The present invention will be described in terms of processes for producing components containing ceramic-based materials, including homogeneous ceramic materials and CMC materials that may contain discontinuous and/or continuous fiber reinforcement materials. While various applications are foreseeable and possible, applications of particular interest include are high temperature applications, for example, components of gas turbines, including land-based and aircraft gas turbine engines. Furthermore, specific reference will be made to airfoil components, including turbine blades and vanes for use within the turbine sections of a gas turbine engine. While the invention is applicable to a wide variety of ceramic-based materials, ceramic-based materials of particular interest to the invention are believed to be CMC materials containing silicon, such as CMC's containing silicon carbide as the reinforcement and/or matrix material, for example, continuous silicon carbide fibers in a matrix of silicon carbide. However, other ceramic-based materials are also within the scope of the invention.
Similar to what was described for the blade 10 of
The airfoil portion 38 of the blade 30 is an excellent candidate for being produced from a ceramic-based material, and especially a CMC material, because it is directly exposed to the hot combustion gases and has a generally linear geometry. On the other hand, the platform portion 36, dovetail portion 32 and its tangs 34 have more complex geometries than the airfoil portion 38, in the sense that the airfoil portion 38 has a generally linear geometry along its dominant axis, whereas the dovetail and platform portions 32 and 36 define geometric features oriented transverse to each of their dominant axes. Furthermore, these off-axis geometric features are detailed interface/supporting features of the blade 30, and therefore require structural interface capabilities that pose substantial challenges to designing, manufacturing and integrating a completely CMC blade (such as the blade 10 of
As used herein, the term shank nub refers to a limited portion, preferably an interior region, of the entire shank portion 40, which further includes the dovetail portion 32 and its tangs 34. As represented in
As a ceramic-based material, the unitary airfoil portion 38 and shank nub 48 can be produced by known ceramic processes. For example, the unitary airfoil portion 38 and shank nub 48 can be CMC materials fabricated from prepregs. Nonlimiting examples include the processes disclosed in U.S. Pat. Nos. 5,015,540, 5,330,854, 5,336,350, 5,628,938, 6,024,898, 6,258,737, 6,403,158, and 6,503,441, and U.S. Patent Application Publication No. 2004/0067316. As a particular example, the unitary airfoil portion 38 and shank nub 48 can be fabricated by the previously-described prepreg melt-infiltration (MI) process, wherein multiple prepregs are formed to contain the desired reinforcement material and a precursor of the CMC matrix material, as well as one or more binders and, depending on the particular desired CMC material, possibly carbon or a carbon source. The prepregs undergo lay-up, are debulked and cured while subjected to elevated pressures and temperatures, and subjected to any other suitable processing steps to form a laminate preform. Thereafter, the laminate preform may be heated in a vacuum or an inert atmosphere to decompose the binders and produce a porous preform that is then melt infiltrated. If the CMC material contains a silicon carbide reinforcement material in a ceramic matrix of silicon carbide (a SiC/SiC CMC material), molten silicon is typically used to infiltrate into the porosity, react with a carbon constituent (carbon, carbon source, or carbon char) within the matrix to form silicon carbide, and fill the porosity. However, it will be apparent from the following discussion that the invention also applies to other types and combinations of CMC materials.
Because of the generally linear geometry of the airfoil portion 38 and shank nub 48, the required lay-up process is not nearly as labor intensive and does not require the level of skilled labor that would be required if the entire blade 30 were to be fabricated from prepregs.
Though the drive for additional turbine engine performance has prompted the desire for using CMC materials due to increased gas path temperatures, those regions of blades (and other turbine components) that are not directly exposed to the hot combustion gases, including the dovetail, platform and shank portions 32, 36 and 40 of the blade 30, may utilize materials with lower temperature capabilities, for example, nickel-, cobalt- or iron-based alloys currently available and used in turbomachinery. Notable but nonlimiting examples include such superalloys as IN (Inconel) 718, René N5 (U.S. Pat. No. 6,074,602), René N6 (U.S. Pat. No. 5,455,120), GTD-444®, René 77 (U.S. Pat. No. 3,457,066), René 80, René 80H and René 125.
As evident from the CMC subcomponent 44 seen in
In view of the above, a metallic material can be cast around the shank nub 48 of the simplified, de-featured CMC subcomponent 44, to produce the entire unitary metallic subcomponent 46, which in effect is an overlaying metal casing that defines the dovetail portion 32 and its tangs 34, as well as what will be referred to as a shank casing 50 that encases the portion of the shank nub 48 above the dovetail portion 32 of the subcomponent 46. In
Because the coefficient of thermal expansion (CTE) of metallic materials that can be used to form the cast metallic dovetail portion 36 and shank casing 50 is typically higher than the CTE of typical CMC materials, during solidification of the metallic material around the CMC shank nub 48 the cast metallic material that defines the cavity 52 will contract more than the CMC material and compress the shank nub 48, providing a compression fit and tight encapsulation of the CMC nub 48 and retention of the CMC subcomponent 44 to the metallic subcomponent 46, which is in addition to the retention capability provided as a result of the subcomponent 46 surrounding the enlarged base 49 of the shank nub 48. As a nonlimiting example, a suitable compression fit is believed to be achievable with a metallic material such as the aforementioned nickel-based superalloy René 80H, which has a CTE of about 14 ppm/° C., in comparison to a CTE of about 4 ppm/° C. for SiC—SiC CMC materials. This CTE differential is capable of yielding a strain of about 1% when cooled to room temperature from a casting temperature of about 2200° F. (about 1200° C.), resulting in a room temperature stress state in which the CMC shank nub 48 is in compression and the metallic subcomponent 46 surrounding the nub 48 is in tension. Particularly for blades whose dovetails are in compression during operation, the shrink-fit resulting from the casting process is capable of providing a robust mechanical attachment.
The process of “co-casting” the metallic subcomponent 46 on the CMC subcomponent 44 can be achieved in a variety of ways.
Other methods that can be used to form the metallic subcomponent 46 include spin casting techniques. As known in the art, spin casting processes are similar to conventional investment casting processes in the fact that a mold is created by coating a wax replica of the part in ceramic, and then removing the wax to yield a female form the part (“mold”), which is then filled with molten metal that solidifies to form the final part. Spin casting techniques depart from conventional casting methods in that the latter relies on gravitational force to act on a molten metal to fill the mold, whereas the mold in a spin casting process is rotated to induce centrifugal forces that act on the molten metal. This additional force is beneficial to certain casting geometries and/or materials to ensure a complete fill of the mold with acceptable microstructure and lack of internal defects. Spin casting also differs from centrifugal casting processes, in which a molten metal is poured from a crucible into a central pour cup that is aligned with the rotational axis of a rotating mold. The molten metal initially has zero centrifugal force acting upon it, and takes a finite amount of time until it flows away from the center of rotation and slowly picks up centrifugal force. With spin casting, the charge (unmelted raw material) is melted at a distance way from the center of rotation, such that when the charge is melted and rotation starts, the molten metal is immediately acted upon by centrifugal force, resulting in a more rapid fill of the mold than either conventional or centrifugal casting processes.
It should be noted here that the subcomponent 46 depicted in
In addition to or as an alternative to relying on the casting technique to minimize undesirable chemical reactions between the CMC material of the subcomponent 44 and the molten metal material of the subcomponent 46, the present invention also contemplates the use of interface coatings provided between the CMC and metallic subcomponents 44 and 46. In addition or alternatively, an interface coating can be employed to enhance thermal expansion compliance for the shrinkage of the metal around the CMC subcomponent 44 during solidification to reduce the incidence of cold cracking.
As evident from
While the invention has been described in terms of specific embodiments, it is apparent that other forms could be adopted by one skilled in the art. Therefore, the scope of the invention is to be limited only by the following claims.
Darkins, Jr., Toby George, Heyward, John Peter, Jamison, Joshua Brian, Estill, Eric Alan, Hawkins, James Thomas, Deines, James Herbert, Marusko, Mark Willard
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